CN103119247B - Gas-turbine blade - Google Patents
Gas-turbine blade Download PDFInfo
- Publication number
- CN103119247B CN103119247B CN201180030863.8A CN201180030863A CN103119247B CN 103119247 B CN103119247 B CN 103119247B CN 201180030863 A CN201180030863 A CN 201180030863A CN 103119247 B CN103119247 B CN 103119247B
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- Prior art keywords
- trailing edge
- passage
- root
- gas
- turbine blade
- Prior art date
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Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/187—Convection cooling
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/10—Stators
- F05D2240/12—Fluid guiding means, e.g. vanes
- F05D2240/122—Fluid guiding means, e.g. vanes related to the trailing edge of a stator vane
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/20—Rotors
- F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
- F05D2240/304—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the trailing edge of a rotor blade
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2250/00—Geometry
- F05D2250/10—Two-dimensional
- F05D2250/18—Two-dimensional patterned
- F05D2250/185—Two-dimensional patterned serpentine-like
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
- F05D2260/205—Cooling fluid recirculation, i.e. after cooling one or more components is the cooling fluid recovered and used elsewhere for other purposes
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
- F05D2260/221—Improvement of heat transfer
- F05D2260/2214—Improvement of heat transfer by increasing the heat transfer surface
- F05D2260/22141—Improvement of heat transfer by increasing the heat transfer surface using fins or ribs
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The present invention relates to gas-turbine blade (2), it comprises: root (6) and aerofoil profile (4), and this aerofoil profile has leading edge (16) and trailing edge (18); Cooling air channels system (28), it extends to the trailing edge passage (56) of trailing edge (18) via roundabout serpentine channel (50) from the cold air openings (44) in root (6), comprise air outlet slit (62) at trailing edge (18) place.In order to the trailing edge (18) of cooled blade (2) effectively, propose this cooling air channels system (28) and comprise and walk around the air bypass passage (54) that the described cold air openings (44) in root (6) is connected to trailing edge passage (56) by serpentine channel (50).
Description
Technical field
The present invention relates generally to turbine blade, and relates more specifically to gas-turbine blade, and it comprises: root; And there is the aerofoil profile of leading edge and trailing edge; Cooling air channels system, it extends to the trailing edge passage of trailing edge via roundabout serpentine channel from the cold air openings in root, it comprises air outlet slit at trailing edge place.
Background technique
Gas turbine is in hot operation, and this high temperature may reach 1200 degrees Celsius and higher.Therefore, turbine blade must can bear such high temperature.In order to extend the life-span of blade, it usually comprises cooling system and carrys out guiding cooling air and pass through blade.
Gas-turbine blade comprises root, platform and from the outward extending aerofoil profile of platform, and aerofoil profile comprises tip, leading edge and trailing edge.Between gas turbine on-stream period, can produce heavily stressed in some regions of turbine blade.In the trailing region at aerofoil profile hub area and hub place, there is specific longevity district, its downstream side in aerofoil profile forms the wall of relative thin.Due to structure and during operation heavily stressed of its relative thin, trailing edge very easily forms crackle, and this can cause aerofoil profile to lose efficacy.
Cooling system comprises internal cooling channel, and its compressor admission of air from gas turbine also makes air pass through blade.Cooling channel comprises multiple stream, and it is designed to turbine blade to remain on relatively uniform temperature.But, the air flowing at centrifugal force and boundary layer place sometimes can stop some district of appropriate cooling turbine bucket, thus causes forming hot localised points, and this can reduce the working life of turbine blade.
Cooling system in aerofoil profile can comprise path of cool air to maximize the convection current cooling in airfoil tip and trailing edge, and is discharged by the Cooling Holes in the tip of aerofoil profile and trailing edge by a part of cooling-air.Such turbine blade is known, such as, from U. S. Patent 4278400.
Summary of the invention
Target of the present invention is to provide the gas-turbine blade at the rear edge part of aerofoil profile with high cooling capacity.
Realize this target according to the present invention by gas-turbine blade as above, wherein cooling air channels system comprises and walks around the air bypass passage that the described cold air openings in root is connected to trailing edge passage by serpentine channel.
The operation of turbogenerator causes producing in multiple regions of turbine blade heavily stressed.The high stress areas that existence one is special in airfoil trailing edge, this airfoil trailing edge is a part for the aerofoil profile forming relatively thin edges.Because trailing edge is relatively thin and be easy to heavily stressed region occurs at run duration, so trailing edge is very easily subject to the impact that crackle is formed, this can cause the inefficacy of aerofoil profile.Utilize bypass channel, the cool air from the opening in root is guided directly into trailing edge passage and can not be heated in radial passage or serpentine channel, thus effectively cools trailing edge.
Cooling-air is supplied to the inside of aerofoil profile via the opening of the radially inner side of root.Such root can have more than one opening.As known, an opening can supply air to serpentine channel and be fed to trailing edge passage further, and another opening can be used as bypass that air is directly fed to trailing edge passage.But if one in described opening for supplying air to front edge area, so only the surplus next one can be used for the opening supplying air to trailing region or passage.The present invention proposes especially for the advantageous schedule of blade, it has more than one opening, especially only two openings, and wherein only an opening supplies air to trailing edge.This opening for supplying serpentine channel and bypass channel, the outer wall of cooled blade and trailing edge effectively.
The trailing edge that trailing edge passage can be parallel to blade extends, and directly leads to trailing edge or around the one or more outlets in the region of trailing edge.
According to an aspect of the present invention, the narrowest width of bypass be aerofoil profile chord width and thus at least 10% of distance between leading edge 16 and trailing edge 18 (especially near the chord width of platform at top forming root).When from the same opening in root to bypass channel and serpentine channel supply, the fact that should be noted that is: enough cooling-airs are supplied via bypass channel.Bypass channel therefore should be comparatively large on hydraulic diameter, especially between 10% and 15% of the chord width of aerofoil profile.This width can be define bypass channel wall between distance, especially extending in the plane of trailing edge from leading edge.
Based on same reasons, if the narrowest width of bypass is greater than the half of the width of the cooling channel separating bypass channel, so this is favourable.
According to a further aspect in the invention, cooling air channels system comprises the root passage being positioned at root at least in part, and bypass channel separates from root passage within root.Owing to being quite low to the heating of the air within root, therefore this embodiment guarantee bypass air after arrival edge passage time be cold.Root passage can extend to further downstream radial passage from the opening in root compared to bypass branch.
Partly be positioned within root if bypass channel is configured at least one of its length, especially at the radially inner side of the platform of blade, then the heating of the blade in bypass channel region be held lower, ensure that the effective cooling to trailing edge.
In another embodiment of the invention, pedestal is arranged in bypass channel, and the cooling-air that described pedestal is through bypass channel surrounds.Heat from region to be cooled can be effectively transferred to cooling-air.The pressure sidewall of aerofoil profile can be connected to suction sidewall or only can be connected to one of wall be projected in bypass channel by pedestal.
If the trailing edge passage that bypass is led at least comprises pedestal in the region of bypass opening, then the continuous air flow that can obtain for Homogeneous cooling trailing edge moves.
Based on same reasons, the preferably quantity of the pedestal of per unit area, i.e. blocking effect is identical in bypass channel with trailing edge passage.And the pedestal in the pedestal preferably in bypass channel and trailing edge passage is identical type.Especially, described pedestal has same shape and size.
If bypass directly leads to trailing edge, the effective cooling to trailing edge can be realized.Following situation is such just for this: when the distance between the opening of bypass channel entering the nearest air outlet slit in trailing edge passage arrival trailing edge and/or the opening of trailing edge passage is less than three times of the narrowest width trailing edge being connected to the bypass channel in the plane of leading edge.
In further embodiment of the present invention, bypass is leading to trailing edge passage in the radial direction from root to tip.Due to the rotation of blade, radial force acts on and flows through on the cooling-air of bypass channel.Described flowing is maintained by radial opening, ensure that the abundant flowing of the cooling-air by bypass channel.
Accompanying drawing explanation
Although require that claim of the present invention is to summarize specification specifically to indicate and to know, but describe embodiments of the invention referring now to accompanying drawing by means of only way of example, in accompanying drawing:
Fig. 1 shows the stereogram of the turbine blade comprising root and aerofoil profile,
Fig. 2 shows to be had for the cross-sectional view of guiding cooling air by the turbine blade of the passage of aerofoil profile, and
Fig. 3 shows the plan view of airfoil tip.
Embodiment
With reference to figure 1, show the exemplary turbine blade 2 for gas turbine engine.Blade 2 comprises aerofoil profile 4 and root 6, and this root 6, for routinely blade 2 being fixed to the rotor disk of motor to be supported in the working media flow path of turbine by blade 2, applies motive force in its surface at this working medium gas.With reference to figure 1 and Fig. 2, aerofoil profile 4 has the outer wall 8 around empty internal 14.Suction sidewall 12(Fig. 3 that aerofoil profile outer wall 8 comprises substantially recessed pressure sidewall 10 and substantially protrudes), described pressure sidewall 10 and suction sidewall 12 broad ways are separated thus are limited empty internal 14 betwixt.Pressure sidewall and suction sidewall 10,12 extend and together with being attached at downstream trailing edge 18 place in upstream leading edge 16 between upstream leading edge 16 and downstream trailing edge 18.Leading edge 16 and trailing edge 18 are axially or to be tangentially separated from each other.The longitudinal direction of the blade 2 that aerofoil profile 4 limits along the span by aerofoil profile 4 or radial direction radial outer leafs tip end surface 22 radial direction from aerofoil profile platform 20 in radial to the tip 24 of aerofoil profile 4 extends.
As shown in Figure 2, in empty internal 14, two cooling channels systems 26,28 are limited.Cooling channels system 26,28 extends through turbine blade 2 and the two equal fluid is communicated in cooling fluid source and is separated from each other along spanwise.Cooling channels system 26,28 is all through aerofoil profile 4 and along its whole length extension heat is delivered to cooling fluid from the surface of aerofoil profile sidewall 10,12 and keeps the temperature of blade 2 lower than maximum allowable temperature pressure sidewall 10 and suction sidewall 12.
Cooling channels system 26 comprises radial passage 30 and follows the axial passage 32 of radial passage 30 along air-flow direction closely.Cooling channels system 26 directly extends along leading edge 16 from the opening 34 of the radial inner end of the root 6 inside outer wall 8, be directly adjacent to leading edge 16 from the leading edge 16 footpath inside beginning until most advanced and sophisticated bottom 36 thus form the wall paralleled with the extension at tip 24.Run through this path, channel system 26 does not have branch thus its cooling-air is all fed to most advanced and sophisticated bottom 36 along leading edge 16 and effectively cools leading edge 16.
Along its further route, cooling channels system 26, or more accurately, its axial passage 32 ends at multiple air outlet slit 38,40,42, and all these outlets are all arranged on tip 24 place of aerofoil profile 4.Therefore, all cooling-airs entering cooling channels system 26 through inside opening 34 are all directed into the outlet 38,40,42 at most advanced and sophisticated 24 top places.
Second cooling channels system 28 also starts from the opening 44 in the radial inner end of the root 6 of blade 2 and extends to most advanced and sophisticated 24 along spanwise.But, this system 28 is branched into multiple passage: two parallel radial passage 46,48, serpentine flow passage 50, tip channel 52, bypass passageways 54 and trailing edge passages 56.Radial passage 46 is parallel to leading edge passage 30 and extends and lead to tip channel 52 and serpentine flow passage 50.The radial wall 58 that radial passage 48 is blocked is separated in radial passage 46, be also parallel to leading edge passage 30 extends and leads to tip channel 52 and serpentine flow passage 50.
Serpentine flow passage 50 starts from the end of radial passage 46,48, to extend into radially-inwardly and radially outward again, and lead to trailing edge passage 56 with two u turns from radially outward direction.Guide radially-inwardly u turn by u turn wall 60, this wall 60 define u turn and with at least 150 degree angles from radially-inwardly changing into radially outward.Trailing edge passage 56 can end at the multiple outlets be arranged in trailing edge 18, and the specific embodiment wherein shown in Fig. 1 and Fig. 2 only comprises a back outlet 62, and it is formed radial fissure and 80% of the radial length of extend through trailing edge 18.Trailing edge passage 56 is shaped to resemble the radial passage opened to the trailing edge in outlet (correspondingly exporting 62) along its axial side.
The root passage 64 extending to radial passage 46,48 from opening 44 is directly connected in trailing edge passage 56 by bypass passageways 54, thus cooling-air is directly directed to trailing edge passage 56 from root passage 64.Bypass passageways 54 bends from root passage 64 at it during the route of trailing edge passage 56, thus radially the section being located immediately at the outlet crack place of trailing edge 18 of trailing edge passage 56 is led in direction, therefore directly leads to trailing edge 18 and correspondingly leads to trailing edge air outlet slit 62.
Root passage 64 is positioned at the root 6 of blade 2 completely, therefore lower than platform 20, is namely positioned at the inner radial of platform 20.Bypass passageways 64 is oriented at least half (particularly more than 3/4 of its length) of its length lower than platform 20.
In order to the well-off cool air of trailing edge passage 56, the most narrow passage width 66 of bypass passageways 54 is greater than the half of the width of the root passage 64 separating bypass passageways 54.This narrowest width is about 11% of the chord width (thus the length between leading edge 16 and trailing edge 18) of aerofoil profile.In this narrowest part of bypass passageways 54, it is perpendicular to the width of channel width 66, that is along from suction sidewall 14 to the width in the direction of pressure sidewall 10, be greater than bypass passageways 54 and enter in the open area of trailing edge passage 56 along from suction sidewall 14 to the width in the direction of pressure sidewall 10 at it.
Inner at trailing edge passage 56, multiple pedestal 68 be oriented to by the cooling-air institute of being flow through trailing edge 56 around.Pedestal 68 is formed circular important actor, its pressure sidewall 10 is connected to suction sidewall 12 and by outer wall 8 generate delivered heat in trailing edge passage 56.The pedestal 68 of identical type is positioned at the downstream section of serpentine channel 50 and bypass passageways 54, this downstream section extends about 2/3 of the total length of bypass passageways 54, and wherein the quantity of the pedestal 68 of per unit area can be identical in bypass passageways 54 and trailing edge passage 56.
Two cooling air channels systems 26,28 are that cooling-air is supplied in the outlet 38,40,42,70 in most advanced and sophisticated 24, but channel system 26 only supply at least one air outlet slit 62 that outlet 38,40,42 in most advanced and sophisticated 24 and channel system 28 supply the trailing edge place of at least one air outlet slit 70 in tip 24 and aerofoil profile 4.The setting of the air outlet slit 38,40,42,70 in Fig. 3 in best visible cusps 24.
Fig. 3 shows the tip 24 of aerofoil profile 2 with plan view.Most advanced and sophisticated 24 comprise rib 72 or outstanding wall, its formed outer wall 8 outermost radial outside section, fully extend around the bottom 36 of most advanced and sophisticated 24 and be preferably so that the 2%-3% lifting of the 1%-2% of the length of blade 2 or the length of aerofoil profile 4 is on bottom 36.Bottom 36 comprises outlet 38,40 and dust export 74, and outlet 38 forms first group, and exports 40 formation second group.In the leading edge 16 that first group of outlet 38 is arranged on most advanced and sophisticated 24 and in leading edge section 76, be referred to as the leading edge at the top of most advanced and sophisticated 24 for simplicity.This section 76 extends to the imaginary line shown in Fig. 3 from leading edge 16, and this imaginary line is vertical with the skeleton line 80 of blade 2 and through the upstream face of aerofoil profile 4 or pressure side surface 10.In the embodiment shown in fig. 3, this section 76 extends certain distance towards trailing edge 18, and this distance is 1/10 of the distance between leading edge 16 and trailing edge 18.Second group of outlet 40 is arranged on the on the pressure side section 78 at tip 24, and be referred to as the top of most advanced and sophisticated 24 for simplicity on the pressure side, it extends to skeleton line 80 from pressure sidewall 10.Two groups of outlets 38,40 supply by the first cooling air channels system 26.
First group of outlet 38 is formed by three holes being all arranged to direct adjacent ribs 72 in bottom 36.Form second group of outlet 40 by five holes being also all arranged to direct adjacent ribs 72 in bottom 36, but distance between described hole to export in 38 than first group wide.The Kong Jun of first group has the same diameter less than the diameter in the hole of second group.Outlet 40 is unequal apart from distance each other.Central exit 40 is greater than the distance of outermost outlet 40 apart from its adjacent outlets 40 of this group apart from the distance of its adjacent outlets 40.
Extend to the No way out district of second group from first group between two groups of outlets 38,40.This district is along observing from leading edge 16 to the direction of trailing edge 18 diameter that is greater than first group of outlet 38 and being greater than the longest distance the hole of second group of outlet 40.
Extend to from trailing edge 18 as shown in Figure 3 imaginary line approximately extend apart from leading edge 16 distance 30% tip 24 trailing edge section 82(be called the trailing edge at the top of most advanced and sophisticated 24 for simplicity), outlet 42,70 is set.They are formed groove or crack, and it is directly defined by rib 72 or outstanding wall and is radially directed towards outer and is inclined to towards trailing edge 18 and becomes about 70 degree with radial direction, wherein 0 degree be absolute radial and 90 degree be bottom being parallel to.Due to this inclination, so two outlets 42,70 are defined by wall radial direction.Outlet 42 is defined by bottom 36 and the wall 84 that the first cooling passage system 26 is located away from the second cooling passage system 28.Outlet 70 is defined by wall 84 and the wall 86 of the trailing edge ends leading to rib 72.
Claims (11)
1. a gas-turbine blade (2), comprising: root (6) and aerofoil profile (4), and this aerofoil profile (4) has leading edge (16), trailing edge (18) and tip (24); Cooling air channels system (28), extend to the trailing edge passage (56) of trailing edge (18) via roundabout serpentine channel (50) from the cold air openings (44) in root (6), air outlet slit (62) is comprised at trailing edge (18) place, and walk around the air bypass passage (54) that the described cold air openings (44) in root (6) is directly connected in trailing edge passage (56) by serpentine channel (50)
It is characterized in that, cooling air channels system (28) comprises the radial passage (46) led to serpentine flow passage (50) and lead to tip channel (52), described cold air openings (44) in root (6) to be connected at least one air outlet slit (70) in the trailing edge section (82) of the radial outer leafs tip end surface (22) on the top of tip (24) by this tip channel (52) via described radial passage (46), thus be described tip (24) top on radial outer leafs tip end surface (22) trailing edge section (82) supply cooling-air.
2. gas-turbine blade according to claim 1 (2),
Be characterised in that the narrowest width of bypass channel (54) is at least 10% of the chord width of aerofoil profile (4).
3. gas-turbine blade according to claim 1 and 2 (2),
Be characterised in that the narrowest width of bypass channel (54) is greater than the half of the width of the root passage (64) separating bypass channel.
4. gas-turbine blade according to claim 1 and 2 (2),
Be characterised in that, cooling air channels system (28) comprises the root passage (64) being positioned at root (6) at least in part, and bypass channel (54) separates from root passage (64) within root (6).
5. gas-turbine blade according to claim 1 and 2 (2),
Be characterised in that bypass channel (54) is configured at least one of its length and is partly positioned within root (6).
6. gas-turbine blade according to claim 1 and 2 (2),
Be characterised in that the pedestal (68) that cooled air stream surrounds is arranged in bypass channel (54).
7. gas-turbine blade according to claim 1 and 2 (2),
Be characterised in that the trailing edge passage (56) that bypass channel (54) leads at least comprises pedestal (68) in the region of bypass opening.
8. gas-turbine blade according to claim 1 and 2 (2),
The quantity of the pedestal (68) of per unit area is identical in bypass channel (54) with trailing edge passage (56).
9. gas-turbine blade according to claim 1 and 2 (2),
Pedestal (68) in bypass channel (54) and the pedestal (68) in trailing edge passage (56) have same shape and size.
10. gas-turbine blade according to claim 1 and 2 (2),
Be characterised in that bypass channel (54) directly leads to trailing edge (18).
11. gas-turbine blades according to claim 1 and 2 (2),
Be characterised in that bypass channel (54) is leading to trailing edge passage (56) in the radial direction from the root (6) of aerofoil profile to most advanced and sophisticated (24).
Applications Claiming Priority (3)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| RU2010000352 | 2010-06-23 | ||
| RUPCT/RU2010/000352 | 2010-06-23 | ||
| PCT/EP2011/060500 WO2011161188A1 (en) | 2010-06-23 | 2011-06-22 | Gas turbine blade |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN103119247A CN103119247A (en) | 2013-05-22 |
| CN103119247B true CN103119247B (en) | 2015-11-25 |
Family
ID=44627761
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201180030863.8A Active CN103119247B (en) | 2010-06-23 | 2011-06-22 | Gas-turbine blade |
Country Status (4)
| Country | Link |
|---|---|
| US (1) | US8702391B2 (en) |
| EP (1) | EP2564029B1 (en) |
| CN (1) | CN103119247B (en) |
| WO (1) | WO2011161188A1 (en) |
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| EP3123000B1 (en) * | 2014-03-27 | 2019-02-06 | Siemens Aktiengesellschaft | Blade for a gas turbine and method of cooling the blade |
| FR3021697B1 (en) * | 2014-05-28 | 2021-09-17 | Snecma | OPTIMIZED COOLING TURBINE BLADE |
| US10294799B2 (en) * | 2014-11-12 | 2019-05-21 | United Technologies Corporation | Partial tip flag |
| US10774654B2 (en) | 2015-07-31 | 2020-09-15 | General Electric Company | Cooling arrangements in turbine blades |
| US10502067B2 (en) | 2016-01-22 | 2019-12-10 | United Technologies Corporation | Dual-fed airfoil tip |
| US10450874B2 (en) * | 2016-02-13 | 2019-10-22 | General Electric Company | Airfoil for a gas turbine engine |
| US10487660B2 (en) * | 2016-12-19 | 2019-11-26 | General Electric Company | Additively manufactured blade extension with internal features |
| US10508548B2 (en) * | 2017-04-07 | 2019-12-17 | General Electric Company | Turbine engine with a platform cooling circuit |
| US10697301B2 (en) * | 2017-04-07 | 2020-06-30 | General Electric Company | Turbine engine airfoil having a cooling circuit |
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| US10519782B2 (en) * | 2017-06-04 | 2019-12-31 | United Technologies Corporation | Airfoil having serpentine core resupply flow control |
| US10612394B2 (en) * | 2017-07-21 | 2020-04-07 | United Technologies Corporation | Airfoil having serpentine core resupply flow control |
| US10837291B2 (en) | 2017-11-17 | 2020-11-17 | General Electric Company | Turbine engine with component having a cooled tip |
| US10655476B2 (en) * | 2017-12-14 | 2020-05-19 | Honeywell International Inc. | Gas turbine engines with airfoils having improved dust tolerance |
| KR102114681B1 (en) * | 2018-09-21 | 2020-05-25 | 두산중공업 주식회사 | Turbine blade having pin-fin array |
| KR102180395B1 (en) | 2019-06-10 | 2020-11-18 | 두산중공업 주식회사 | Airfoil and gas turbine comprising it |
| CN110410158B (en) * | 2019-08-16 | 2022-04-12 | 杭州汽轮动力集团有限公司 | Turbine rotor blade of gas turbine |
| US11136890B1 (en) * | 2020-03-25 | 2021-10-05 | General Electric Company | Cooling circuit for a turbomachine component |
| CN112160796B (en) * | 2020-09-03 | 2022-09-09 | 哈尔滨工业大学 | Turbine blade of gas turbine engine and control method thereof |
| US12006836B2 (en) * | 2021-07-02 | 2024-06-11 | Rtx Corporation | Cooling arrangement for gas turbine engine component |
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- 2011-06-22 US US13/806,204 patent/US8702391B2/en active Active
- 2011-06-22 WO PCT/EP2011/060500 patent/WO2011161188A1/en active Application Filing
- 2011-06-22 EP EP11728818.3A patent/EP2564029B1/en active Active
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Also Published As
| Publication number | Publication date |
|---|---|
| US8702391B2 (en) | 2014-04-22 |
| EP2564029B1 (en) | 2014-10-01 |
| US20130209268A1 (en) | 2013-08-15 |
| EP2564029A1 (en) | 2013-03-06 |
| WO2011161188A1 (en) | 2011-12-29 |
| CN103119247A (en) | 2013-05-22 |
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